Combustor for gas turbine engine and gas turbine

ABSTRACT

A combustor for gas turbine engine, wherein: a plurality of impingement-cooling holes are opened passing through an external wall of a liner for cooling air to blow out towards the outer surface of an internal wall of the liner; a plurality of pin-fins are formed on the outer surface of the internal wall of the liner; a plurality of effusion cooling holes are opened passing through the internal wall of the liner for cooling air to blow out along the inner surface of the internal wall of the liner. The top surface of each pin-fin is not in contact with the inner surface of the external wall of the liner and the ratio of the height of the pin-fin to the equivalent diameter of the impingement-cooling hole is set to be 1.0-3.0.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of InternationalApplication No. PCT/JP2012/058330, filed on Mar. 29, 2012, which claimspriority to Japanese Patent Application No. 2011-078486, filed on Mar.31, 2011, the entire contents of which are incorporated by referencesherein.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to: a combustor for a gas turbine enginewhich is used in gas turbine engines such as an aircraft gas turbineengine and a power generation gas turbine engine, and which isconfigured to produce a combustion gas by combusting fuel in compressedair; and a gas turbine.

2. Description of the Related Art

A general annular combustor for a gas turbine engine includes an annularcombustor case and an annular combustor liner is concentricallyinstalled inside the combustor case. As a result, an annular combustionchamber for combusting fuel is formed inside the combustor liner. Inaddition, fuel nozzles configured to inject the fuel into the combustionchamber are installed in a front section of the combustor liner atintervals in the circumferential direction; and swirlers configured tointroduce compressed air into the combustion chamber are installedaround the peripheral edge of each fuel nozzle in the front section ofthe combustor liner.

On the other hand, a combustor for a gas turbine engine which enablesboth effusion cooling and impingement cooling to be applied to thecombustor liner has been developed for the purpose of enhancing thecooling performance of the combustor liner (see JP 2010-43643 A). To putit specifically, the combustor liner in the combustor for a gas turbineengine of the prior art has a double-wall structure including anexternal wall of the liner and an internal wall of the liner, andmultiple impingement-cooling holes are opened through the external wallof the liner for part of compressed air as cooling air to blow outtowards the outer surface (outer wall surface) of the internal wall ofthe liner. In addition, multiple effusion cooling holes are openedthrough the internal wall of the liner for the cooling air to blow outalong the inner surface (inner wall surface) of the internal wall of theliner.

SUMMARY OF THE INVENTION

Incidentally, there is a growing demand in recent years for higher powerof a gas turbine engine. To meet the demand, the turbine inlettemperature, namely, the combustion gas temperature inside a combustionchamber tends to become very high. For this reason, it has becomeimperative that the cooling performance of the combustor liner exposedto the combustion gas should be enhanced to a higher level.

In view of the above, an object of the present invention is to provide acombustor having a novel configuration for a gas turbine engine and toprovide a gas turbine, each of which can enhance the cooling performanceof a combustor liner to a higher level.

For the purpose of solving the foregoing problem, the inventors of thisapplication carried out a cooling performance test as follows. Testpieces simulating a combustor liner having a double-wall structure inwhich, multiple impingement-cooling holes were opened passing through anexternal wall of a liner, multiple effusion cooling holes were openedpassing through an internal wall of the liner, and pin-fins (heattransfer enhancement pin-fins) were formed on the outer surface of theinternal wall of the liner were provided for cooling performance tests.A large number of test pieces with different pin-fin heights wereprepared. Then, the investors performed the cooling performance test oneach of the test pieces while making a high-temperature gas flow on oneside (the internal wall side of the liner) and making cooling gas flowon the other side (the external wall side of the liner) of the testpiece, respectively. As a result of the cooling performance tests on thelarge number of test pieces (see Example to be described later), theinvestors have succeeded in obtaining a novel finding that coolingeffectiveness of the combustor liner can be sufficiently enhanced whileinhibiting an increase in the weight of the combustor liner by setting aratio of the height of the pin-fins to the equivalent diameter of theimpingement-cooling holes to an appropriate ratio with the top surfaceof each pin-fins being not in contact with the inner surface of theexternal wall of the liner. Based on the finding, the inventors havecompleted the present invention. Here, the appropriate ratio is in arange of 1.0 to 3.0. In addition, the inventors consider that theabove-mentioned novel finding resulted from a fully heat transferenhancement by the pin-fins.

A first aspect of the present invention is a combustor for a gas turbineengine used in the gas turbine engine and configured to produce acombustion gas by combusting fuel in compressed air. The combustorcomprises: a combustor case; and a combustor liner provided inside thecombustor case and including a combustion chamber which is formed insidethe combustor liner, the combustion chamber configured to combust thefuel, wherein the combustor liner is formed into a double-wall structureincluding an external wall of the liner and an internal wall of theliner, a plurality of impingement-cooling holes are opened passingthrough the external wall of the liner for part of the compressed air ascooling air to blow out towards an outer surface (outer wall surface) ofthe internal wall of the liner, a plurality of effusion cooling holesare opened passing through the internal wall of the liner for thecooling air to blowout along an inner surface (inner wall surface) ofthe internal wall of the liner, a plurality of pin-fins (heat transferenhancement pin-fins) are formed on the outer surface of the internalwall of the liner, an top surface of each pin-fin is not in contact withan inner surface of the external wall of the liner, and a ratio of aheight of each heat transfer pin to an equivalent diameter of eachimpingement-cooling hole is set in a range of 1.0 to 3.0.

A second aspect of the present invention is a gas turbine comprising thecombustor of the first aspect.

The present invention can sufficiently enhance cooling effectiveness ofthe combustor liner while inhibiting an increase in weight of thecombustor liner. The present invention accordingly makes it possible toenhance a cooling performance of the combustor liner to a higher levelwhile promoting reduction in weight of the combustor for a gas turbineengine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a conceptual diagram of an aircraft gas turbine engine of anembodiment of the present invention.

FIG. 2 is a front cross-sectional view of a combustor for a gas turbineengine of the embodiment of the present invention.

FIG. 3 is an enlarged cross-sectional view taken along the line of FIG.2.

FIG. 4A is a cross-sectional view showing a main part of an externalwall of a liner or an internal wall of the liner

FIG. 4B is a diagram showing a part of a combustor liner in a statedeveloped into a plane.

FIG. 5A is a graph showing a relationship between an effective heattransfer area enhancement ratio and a ratio (H/L) of a height of pin-finto an equivalent diameter of an impingement-cooling hole.

FIG. 5B is a diagram for explaining a cooling performance test.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

Descriptions will be provided for an embodiment of the present inventionby referring to FIG. 1 to FIG. 4B. It should be noted that in thedrawings, reference sign “FF” indicates forward (upstream of themainstream), and reference sign “FR” indicates rearward (downstream ofthe mainstream).

A gas turbine engine (gas turbine) of the embodiment is used, forexample, in an aircraft and a power generator. As shown in FIG. 1, thegas turbine engine (gas turbine) 1 of the embodiment includes acompressor 3. The compressor 3 compresses air A which is taken into thegas turbine engine 1, and thus produces compressed air PA. In addition,a turbine 5 is connected to the compressor 3 by use of a turbine shaft(rotary shaft) 7. The turbine 5 is driven by a combustion gas G whichexpands in the turbine 5, and concurrently drives the compressor 3 byuse of the turbine shaft 7. Furthermore, a combustor 9 is providedbetween the compressor 3 and the turbine 5. The combustor 9 combustsfuel F in the compressed air PA which is sent from the compressor 3,produces the combustion gas G, and discharges the combustion gas Gtowards the turbine 5.

The overall configuration of the combustor 9 will be described asfollows.

As shown in FIG. 2 and FIG. 3, the combustor 9 is, for example, anannular combustor. In this case, the combustor 9 includes an annular(hollow annular) combustor case 11. In addition, the combustor case 11includes: an annular outer case 13; and an annular inner case 15provided inside the outer case 13. Furthermore, an annular introductionport 17 configured to introduce the compressed air PA from thecompressor 3 into the combustor case 11 is provided in the front sectionof the combustor case 11.

An annular (hollow annular) combustor liner 19 is installed in thecombustor case 11. The combustor liner 19 includes: an annular outerliner 21; and an annular inner liner 23 installed inside the outer liner21. In addition, an annular combustion chamber 25 configured to combustthe fuel F is formed in the combustor liner 19. In other words, theannular combustion chamber 25 is formed between the outer liner 21 andthe inner liner 23. It should be noted that the outer case 13, the innercase 15, the outer liner 21, and the inner liner 23 are setconcentrically.

Multiple fuel nozzles 27 configured to inject the fuel F into thecombustion chamber 25 are installed in the front section of thecombustor liner 19 at intervals in the circumferential direction. A fuelpipe 29 configured to supply the fuel F is connected to each fuel nozzle27. Each fuel pipe 29 juts out from the combustor case 11. In addition,a swirler 31 is installed around the peripheral edge of each fuel nozzle27 in the front section of the combustor liner 19. The swirler 31 guidesthe compressed air PA into combustion chamber 25 with adding swirl tothe flow. Furthermore, multiple ignition plugs 33 configured to lightoff (ignite) the fuel F are installed to the combustor case 11. The tipend portion of each ignition plug 33 projects inside the combustor case11 (inside the combustion chamber 25).

The configuration of characteristic portions of the combustor 9 will bedescribed as follows.

As shown in FIG. 3 and FIG. 4A, the outer liner 21 and the inner liner23 each have a double-wall structure which includes an external wall 35of the liner and an internal wall 37 of the liner. Multipleimpingement-cooling holes 39 are opened passing through the externalwall 35 of the liner for part of the compressed air PA to blow outtowards an outer surface (outer wall surface) of the internal wall 37 ofthe liner as cooling air CA. The center line 39 c of eachimpingement-cooling hole 39 is in parallel to the thickness direction TPof the external wall 35 of the liner.

Multiple effusion cooling holes 41 are opened passing through theinternal wall 37 of the liner for the cooling air CA to blow out alongthe inner surface (inner wall surface) of the internal wall 37 of theliner. The center line (center axis) 41 c of each effusion cooling hole41 tilts from the thickness direction TP of the internal wall 37 of theliner such that an outlet portion of the effusion cooling hole 41 islocated downstream of an inlet portion of the effusion cooling hole 41.Here, the diameter of the cross section of each effusion cooling hole 41which is perpendicular to the center line 41 c (or the equivalentdiameter of the effusion cooling hole 41), and the diameter of the crosssection of each impingement-cooling hole 39 which is perpendicular tothe center line (center axis) 39 c (or the equivalent diameter of theimpingement-cooling hole 39) are each set arbitrarily depending on thedesign. It should be noted that the equivalent diameter of a hole means:the diameter of the cross section of the hole when the cross section ofthe hole perpendicular to the center line is a circle; or the hydraulicdiameter (4×area of cross section/peripheral length) when the crosssection thereof is not a circle.

A ratio (Z/D) of a clearance Z between the inner surface of the externalwall 35 of the liner and the outer surface of the internal wall 37 ofthe liner to the equivalent diameter D of each impingement-cooling hole39 is set in a range of 1.0 to 5.0 so that the impingement coolingperformance can be fully exerted.

Multiple pin-fins (heat transfer enhancement pin-fins) 43 are formed onthe outer surface of the internal wall 37 of the liner. A top surface ofeach pin-fin 43 is not in contact with the inner surface of the externalwall 35 of the liner. In addition, a ratio (H/D) of a height H of eachpin-fin 43 to the equivalent diameter D of each impingement-cooling holeis set in a range of 1.0 to 3.0. The reason why the ratio (H/D) is setat 1.0 or greater is that the heat transfer enhancement effect by thepin-fin 43 is not sufficiently exerted if the ratio (H/D) is set below1.0. The reason why the ratio (H/D) is set less than 3.0 is that, evenif the ratio (H/D) is set greater than 3.0, no improvement in the heattransfer by the pin-fin 43 is expected and the weight of the combustorliner 19 is increased. It should be noted that the cross-sectional shapeof each pin-fin 43 is not limited to a square as shown in FIGS. 4A and4B, and may be set in any arbitrary shape.

As shown in FIG. 4B, the outlet portion (the opening on the outlet side)of each impingement-cooling hole 39, the inlet portion (the opening onthe inlet side) of each effusion cooling hole 41, and each pin-fin 43are placed respectively in different locations in a state where thecombustor liner 19 is developed into a plane.

Descriptions will be provided for the work and effect of the embodimentof the present invention on the basis of the above-describedconfiguration.

The compressed air PA sent from the compressor 3 is introduced into thecombustor case 11 through the introduction port 17, and is subsequentlyswirled by the swirler 31 then guided into the combustion chamber 25. Inthe meantime, the fuel F is injected from the multiple fuel nozzles 27into the combustion chamber 25, and the fuel F is ignited by theignition plugs 33. Thereby, the fuel F is combusted in the compressedair PA inside the combustion chamber 25, and the combustion gas G isproduced. The combustion gas G is discharged to the turbine 5, and theoperation of the gas turbine engine 1 is thus continued.

During the operation of the gas turbine engine 1, the compressed air PAflows into an interstice between the inner surface of the combustor case11 and the outer surface of the combustor liner 19. This compressed airPA is caused to blow out towards the outer surface of the internal wall37 of the liner through the multiple impingement-cooling holes 39. Thecooling air CA thus blowing out impinges on the outer surface of theinternal wall 37 of the liner, and performs impingement cooling on theinternal wall 37 of the liner. Furthermore, the cooling air CA havingcontributed to the impingement cooling is caused to blow out along theinner surface of the internal wall 37 of the liner through the multipleeffusion cooling holes 41. Thus, the cooling air CA flows along theinner surface of the internal wall 37 of the liner, and performseffusion cooling on the internal wall 37 of the liner. Moreover, sincethe top surface of each pin-fin 43 is not in contact with the innersurface of the external wall 35 of the liner and the ratio (H/D) of theheight H of each pin-fin 43 to the equivalent diameter D of eachimpingement-cooling hole 39 is set in the range of 1.0 to 3.0, theapplication of the above-mentioned new finding makes it possible tosufficiently enhance the cooling effectiveness of the combustor liner 19while inhibiting an increase in weight of the combustor liner 19 andinhibiting a rise in temperature of the combustor liner 19 (particularlya rise in temperature of the internal wall 37 of the liner).

Accordingly, the embodiment makes it possible to set an optimumstructure in light of the weight and cooling performance of thecombustor 9 for a gas turbine engine.

It should be noted that the present invention is not limited to theabove descriptions of the embodiment and that the present invention canbe carried out in various other modes, inclusive of a mode to apply thetechnical idea employed in the gas turbine engine 1 to a powergeneration gas turbine engine (not illustrated). Furthermore, the scopeof the rights to be encompassed by the present invention is not limitedto these embodiments.

Example

Descriptions will be provided for an example of the present invention byreferring to FIGS. 5A and 5B.

The cooling effectiveness of the combustor liner can be evaluated by aneffective heat transfer area enhancement ratio. The effective heattransfer area enhancement ratio means a ratio of a product of a heattransfer area on a cooling side without the pin-fins and an average heattransfer coefficient to a product of a heat transfer area on the coolingside with the pin-fins and the average heat transfer coefficient. Ahigher effective heat transfer area enhancement ratio leads to highercooling effectiveness.

FIG. 5A shows a relationship between the ratio (H/D) and the effectiveheat transfer area enhancement ratio. Here, reference sign H denotes theheight of the above-described pin-fins and reference sign D denotes theequivalent diameter of the above-described impingement-cooling holes.The relationship shown in FIG. 5A was obtained by correcting ananalytical result obtained by a CFD (Computational Fluid Dynamics)analysis with a result of a cooling performance test. In this coolingperformance test, a test piece 59 (see FIG. 5B) made to simulate thecombustor liner 19 was used, and a surface temperature on an internalwall 77 side of the liner of the test piece 59 was measured in the casewhere a high-temperature gas HG was flowing on the internal wall 77 sideof the liner and the cooling air CA was flowing on an external wall 75side of the liner of the test piece 59. Here, multipleimpingement-cooling holes 79 were opened in the external wall 75 of theliner and multiple effusion cooling holes 81 were opened in the internalwall 77 of the liner. In addition, multiple pin-fins 83 were formed onthe external surface of the internal wall 77 of the liner. The topsurface of each pin-fin 83 was not in contact with the inner surface ofthe external wall 75 of the liner. In other words, theimpingement-cooling holes 79, the effusion cooling holes 81, and thepin-fins 83 in FIG. 5B correspond to the impingement-cooling holes 39,the effusion cooling holes 41, and the heat transfer pins 43 in FIG. 4A,respectively.

It was found from the result of the above-mentioned CFD analysis and thelike that the effective heat transfer area enhancement ratio increasedwhen the ratio (H/D) of the height H of the pin-fins to the equivalentdiameter D of the impingement-cooling holes was in the range of 1.0 to3.0.

In other words, it was found that the cooling performance of thecombustor liner was sufficiently enhanced by setting the ratio (H/D) ofthe height H of the pin-fins to the equivalent diameter D of theimpingement-cooling holes equal to 1.0 or above while making the topsurface of each pin-fin not in contact with the inner surface of theexternal wall of the liner. Furthermore, it was found that the coolingperformance of the combustor liner was not enhanced any more when theratio (H/D) of the height H of each pin-fin to the equivalent diameter Dof the impingement cooling-hole exceeded 3.0.

1: A combustor for a gas turbine engine used in the gas turbine engineand configured to produce a combustion gas by combusting fuel incompressed air, the combustor comprising: a combustor case; and acombustor liner installed inside the combustor case and including acombustion chamber which is formed inside the combustor liner, thecombustion chamber configured to combust the fuel, wherein the combustorliner is formed into a double-wall structure including an external wallof the liner and an internal wall of the liner, a plurality ofimpingement-cooling holes are opened passing through the external wallof the liner for part of the compressed air as cooling air to blow outtowards an outer surface of the internal wall of the liner, a pluralityof effusion cooling holes are opened passing through the internal wallof the liner for the cooling air to blow out along an inner surface ofthe internal wall of the liner, a plurality of pin-fins are formed onthe outer surface of the internal wall of the liner, a top surface ofeach pin-fin is not in contact with an inner surface of the externalwall of the liner, and a ratio of a height of each pin-fin to anequivalent diameter of each impingement-cooling hole is set in a rangeof 1.0 to 3.0. 2: The combustor for a gas turbine engine of claim 1,wherein a ratio of a clearance between the inner surface of the externalwall of the liner and the outer surface of the internal wall of theliner to the equivalent diameter of the impingement-cooling hole is setin a range of 1.0 to 5.0. 3: The combustor for a gas turbine engine ofclaim 1, wherein an outlet portion of each impingement-cooling hole, aninlet portion of each effusion cooling hole, and each pin-fin are placedin mutually different locations in a state where the combustor liner isdeveloped into a plane. 4: The combustor for a gas turbine engine ofclaim 2, wherein an outlet portion of each impingement-cooling hole, aninlet portion of each effusion cooling hole, and each pin-fin are placedin mutually different locations in a state where the combustor liner isdeveloped into a plane. 5: A gas turbine comprising the combustor for agas turbine engine of claim 1.